1. Field of the Invention
This invention relates, in general, to gas turbine engines and, more particularly, to means for providing an annular temperature control air supply to turbine sections, especially turbine nozzles and turbine blades.
2. Description of the Prior Art
A gas turbine engine of the type anticipated in this invention is described in U.S. Pat. Nos. 4,187,054 and 4,214,851, commonly assigned herewith and included herein by reference, includes a fan powered by a low pressure turbine (LPT), a low pressure compressor (LPC), sometimes called a booster, that is also powered by the LPT, a high pressure compressor (HPC) powered by a high pressure turbine (HPT) and a combustor. The combustor is supplied with fuel that is mixed with compressed air from the HPC and ignited to produce hot combustion gas. As the hot combustion gas expands axially out of the gas turbine engine, it impinges first on the HPT and second on the LPT. The HPT transfers some of the combustion energy to the HPC for compressing air used in generating the combustion gas. The LPT extracts some more energy from the combustion gas and uses it to power the fan and the LPC. The fan generates thrust and the LPC provides partially compressed air to the HPC. The remaining energy contained in the combustion gas exits the gas turbine engine and also provides thrust. The fan generally provides most of the thrust.
During normal operation of the gas turbine engine, combustion gas is produced that can reach very high temperatures, typically in excess of 2000.degree. F. which would degrade the strength of the materials, typically metal, used to construct a gas turbine engine if steps were not taken to reduce the material temperature. The present state of the art uses various cooling methods to prevent components from reaching the temperature of the combustion gas. The cooling method anticipated in the present invention extracts air from the HPC and reroutes it past the combustor to the HPT and the LPT sections. Nonrotating blades, called either stator blades or a nozzle, are located between the rotating blades of the HPT and the LPT to efficiently direct the combustion gas to the LPT blades where energy is extracted from the combustion gas. All parts of the HPT and the LPT must be efficiently cooled to prevent material degradation. The need to use the compressed air from the HPC for cooling reduces the efficiency of the gas turbine engine, so it is desirous to provide a cooling system that does not extract more air from the HPC than is necessary to perform the cooling function.
Air transfer tubes, also known as spoolies, are presently used to dispense temperature control air from an annular air supply to an annular turbine nozzle formed of segmented turbine nozzle sections. The temperature control air is supplied from a bleed system connected to a HPC section of the engine. The temperature control air is fed to an annular supply system located around a turbine section and having one side formed, in part, from an annular nozzle support. There are segmented turbine nozzle sections, each having its own manifold that provides temperature control air to the center of the nozzle blades in that section. In order to supply the temperature control air to the manifolds on the segmented nozzle sections, there is at least one air transfer tube that conducts air to the manifold. Each air transfer tube can be interference fit at the manifold and at the annular nozzle support to prevent temperature control air leakage. The air transfer tubes are permitted to slide between slide stops as the turbine nozzle section expands and contracts relative to the annular temperature control air supply manifold. Presently, an air transfer bushing is spot-welded into an aperture in the annular support and has an interior surface where a circumferential groove or key is located to act as a retaining ring seat. A spring-loaded retaining ring engages the key and forms a slide stop to prevent the air transfer tube from traveling past the end of the bushing. A problem arises because multiple temperature cycles between ambient air temperatures and combustion air temperatures of the engine create stress concentrations at the spot welds on the bushing which then act as initiation sites for cracks that propagate to the surrounding structure. The cracks are very difficult and expensive to repair because of their location in the engine. A method to eliminate the welds, to make the air transfer assembly tolerant to the cycling between temperature extremes experienced in a gas turbine engine and to make the bushings replaceable has been devised.
The present invention eliminates the welds and permits the air transfer tube to be replaced. The present invention also provides improved performance and minimizes installation and maintenance costs.
Accordingly, it is an object of this invention to eliminate welds from a gas turbine engine air transfer assembly.
It is a further object of this invention to provide a gas turbine engine having an air transfer assembly that is replaceable.
It is a further object of this invention to provide a gas turbine engine having an air transfer assembly that is easily produced and maintained.